US9869184B2 - Gas turbine blade - Google Patents

Gas turbine blade Download PDF

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Publication number
US9869184B2
US9869184B2 US14/677,343 US201514677343A US9869184B2 US 9869184 B2 US9869184 B2 US 9869184B2 US 201514677343 A US201514677343 A US 201514677343A US 9869184 B2 US9869184 B2 US 9869184B2
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United States
Prior art keywords
leading edge
trailing edge
blade
profile
center point
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Fee Related, expires
Application number
US14/677,343
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English (en)
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US20150285080A1 (en
Inventor
Norbert Huebner
Inga Mahle
Ulrich Haid
Kai Koerber
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MTU Aero Engines AG
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MTU Aero Engines AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
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Assigned to MTU Aero Engines AG reassignment MTU Aero Engines AG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: HAID, ULRICH, MR., KOERBER, KAI, MR., MAHLE, INGA, MS., HUEBNER, NORBERT, MR.
Publication of US20150285080A1 publication Critical patent/US20150285080A1/en
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2240/00—Components
    • F05B2240/10—Stators
    • F05B2240/12—Fluid guiding means, e.g. vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2240/00—Components
    • F05B2240/20—Rotors
    • F05B2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05B2240/301—Cross-section characteristics
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/301—Cross-sectional characteristics
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/712—Shape curved concave
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
    • Y02T50/673—

Definitions

  • the present invention relates to a blade for a gas turbine as well as to a gas turbine having at least one such blade and a method for the dimensional design of such a blade.
  • WO 2012/147938 is a gas turbine blade with an asymmetrical trailing edge, the blade thickness of which is relatively constant extending to the trailing edge.
  • An object of an embodiment of the present invention is to improve a gas turbine.
  • the present invention provides a gas turbine with a corresponding blade and a method for the dimensional design of such a blade.
  • Advantageous embodiments of the invention are set forth herein.
  • a gas turbine blade comprises, in general, an upstream leading edge and a downstream trailing edge, which are connected by a discharge or pressure side, which, in particular, is concave, and an intake side, which, in particular, is convex.
  • a blade element is constructed from profile cross sections, which are arranged so as to overlap along a stacking axis, for diverting the flow, wherein said axis passes through the centers of gravity of the profile cross-sectional surface areas and, in one embodiment of the present invention, is curved or is at least substantially rectilinear in the peripheral and/or axial direction of the gas turbine and can extend in the radial direction of the gas turbine or be inclined toward it.
  • the blade In a cross section perpendicular to the stacking axis or the radial direction, the blade has a profile or an outer contour.
  • This profile has a common pressure-side profile tangent at a leading edge region and at a trailing edge region.
  • Said profile tangent can be formed, in particular, by a line or can be a line, which is placed or rests on the profile from the pressure side, preferably so that it contacts the profile in at least two points, without intersecting it, that is, in graphic terms, without intersecting the line on which the profile would rest on the pressure side.
  • a leading edge region in terms of the present invention extends from the leading edge toward the trailing edge, particularly over at most 25% of a chord length of the profile; a trailing edge region correspondingly extends from the trailing edge toward the leading edge, particularly over at most 25% of the chord length of the profile.
  • the common profile tangent at the leading edge region and at the trailing edge region in terms of the present invention results from the fact that, starting from the leading edge, a point on the profile on the pressure side shifts in the direction of the trailing edge until the tangent at the point on the profile also contacts the profile in the trailing edge section.
  • a tangent tilts at the profile, starting from the leading edge, from the pressure side toward the trailing edge region and then forms the common profile tangent.
  • the profile has a leading edge tangent at the leading edge, which is perpendicular to the common profile tangent. In graphic terms, this is formed or can be formed, in particular, by a normal line to the common profile tangent, which is placed or rests on the profile in the flow direction. In a corresponding way, the profile has a trailing edge tangent at the trailing edge that is parallel to the leading edge tangent and is perpendicular to the common profile tangent and, in graphic terms, is formed or can be formed by a normal line to the common profile tangent, which will be placed or rests on the profile opposite to the flow direction.
  • the profile has a distinct tangent at a contact point with the common profile tangent, at a contact point with the leading edge tangent, and/or at a contact point with the trailing edge tangent, that is, in particular, it can always be differentiated, then the common profile tangent, the leading edge tangent, or the trailing edge tangent can be this distinct or mathematical tangent.
  • the common profile tangent would particularly be a line on which the pressure side of the profile would rest; at the leading edge, the line perpendicular to it would be placed on the profile in the flow direction, and at the trailing edge tangent, the line perpendicular to the common profile tangent would be placed on the profile opposite to the flow direction.
  • tangent is not to be understood in the strictly mathematical sense in the present case, but rather in general terms as defined above as “contacting” or “resting against/resting on.”
  • the circle whose center point lies at the point of intersection of the normal lines or perpendicular lines at the contact point of the trailing edge tangent on the profile with the profile center line, and includes this contact point defines a trailing edge circle in terms of the present invention.
  • the circle that includes at least one point of the pressure side and at least one point of the intake side, and whose center point lies at the point of intersection of the profile center line with the truncated trailing edge and/or which has a minimum diameter defines a trailing edge circle in terms of the present invention.
  • the profile has a camber line, which is defined in terms of the present invention such that it extends, at an equal distance from the pressure side and the intake side, from the center point of the leading edge circle to the center point of the trailing edge circle.
  • the camber line thus forms a well-defined segment of the profile center line, which extends through the profile, at an equal distance from the pressure side and the intake side.
  • the distance between the pressure side and the intake side at a point of the camber line defines a blade thickness at this point of the camber line in terms of the present invention.
  • this can correspond particularly to the diameter of a circle that is inscribed in the profile or contacts the pressure side and the intake side, respectively, in at least one point, which does not intersect the pressure side and the intake side, and the center point of which lies at this point of the camber line.
  • a certain increase in blade thickness in a specific end region of at least one section of the blade element can improve the performance of the blade and, in particular, can reduce any undesired detachment of the flow at the trailing edge.
  • a blade in at least one segment of the stacking axis, has a thickness that, starting from the trailing edge thickness at the center point of the trailing edge circle in a camber line segment, whose length is at least 15%, in particular at least 19%, and at most 25%, preferably at most 21%, of the length of the camber line between the center point of the leading edge circle and the center point of the trailing edge circle, and extending toward the center point of the leading edge circle or in the direction of the center point of the leading edge circle, increases to at least two times and, in particular, to at least 2.5 times the trailing edge thickness.
  • the blade thickness increases in a monotonic manner and, in particular, in a strictly monotonic manner.
  • a (strictly) monotonic increase is understood in the present case in the conventional technical sense to mean that the blade thickness, at each first point of the camber line that, as measured on the camber line, lies closer to the center point of the leading edge circle than an arbitrary second point of the camber line, is at least as large as (in a monotonic manner) and/or always larger (in a strictly monotonic manner) than the blade thickness at the second point.
  • the performance of the blade is further improved; in particular, any undesired detachment of the flow at the trailing edge can be more strongly reduced.
  • the blade thickness, in a region of the camber line segment nearer the trailing edge, this region starting from the center point of the trailing edge circle and the length of which is at most 50% of the length of the camber line segment increases linearly extending toward the center point of the leading edge circle. As a result of such a linear change in the blade thickness only on the trailing edge side, it is possible to improve the transition into the remaining pressure side and intake side.
  • the blade thickness can increase linearly in region of the camber line segment that is nearer the trailing edge, this region starting from the center point of the trailing edge circle and the length of which is at least 50% and, in particular, up to 100% of the length of the camber line segment, toward the center point of the leading edge circle. As a result of such a longer linear change in blade thickness, it is possible to improve the manufacture and/or structural characteristics.
  • a maximum blade thickness is at least 15% and/or at most 30% of the length of the camber line between the center point of the leading edge circle and the center point of the trailing edge circle.
  • this region extending to an end of the camber line segment nearer the leading edge, the length of this region being at least 50% of the length of the camber line segment, the blade thickness increases in a monotonically increasing manner toward the center point of the leading edge circle.
  • the course of the blade thickness along the camber line is concave in the camber line segment or at least in its portion nearer the leading edge.
  • a maximum blade thickness is at least 4% and/or at most 15% of the length of the camber line between the center point of the leading edge circle and the center point of the trailing edge circle.
  • this region extending to an end of the camber line segment nearer the leading edge, the length of this region being at least 50% of the length of the camber line segment, the blade thickness increases in a monotonically decreasing manner toward the center point of the leading edge circle.
  • the course of the blade thickness is convex along the camber line in the camber line segment or at least in its portion nearer the leading edge.
  • a maximum blade thickness is at least 15% and/or at most 30% of the length of the camber line between the center point of the leading edge circle and the center point of the trailing edge circle
  • the blade thickness increases in a region of the camber line segment, this region extending to an end of the camber line segment nearer the leading edge, the length of which is at least 50% of the length of the camber line segment, in a monotonically increasing manner toward the center point of the leading edge circle.
  • the trailing edge thickness is at most 0.5 mm, in particular at most 0.4 mm, preferably at most 0.35 mm. Additionally or alternatively, a wedge angle at the trailing edge is at least 9° and/or at most 17°. These values have been found to be especially advantageous.
  • the segment of the stacking axis in which the profile is designed as explained above, extends over at least 25% and, in particular, at least 75% of a height of the blade element.
  • the profile can be designed, likewise as explained above, so as, in particular, to further improve the aerodynamic performance, or it can be designed differently from this, particularly with lesser blade thickening in the direction to the leading edge, so as to further improve the manufacture and/or thermal and/or mechanical performance.
  • the segment of the stacking axis in which the profile is designed as explained above extends from at least 30% to at least 70% of the height of the blade element or at least in a central region of the blade element.
  • a blade in accordance with the invention is a guide vane that is stationary or fixed in place in the housing, particularly stationary or moving, for a compressor stage or a turbine stage of the gas turbine, or a guide vane of a compressor stage or a turbine stage of a gas turbine in accordance with the invention.
  • a blade in accordance with the invention is mounted so as to rotate or is a rotating blade affixed to the rotor for a compressor stage or a turbine stage of the gas turbine, or a rotating blade of a compressor stage or a turbine stage of a gas turbine in accordance with the invention.
  • a profile in accordance with the invention can improve the performance of guide vanes as well as rotating blades of both compressor and turbine stages.
  • a blade in accordance with the invention is used with particular advantage in an aircraft engine.
  • the profile of a blade in particular of a guide vane or rotating blade of a compressor stage or a turbine stage of a gas turbine, particularly of an aircraft engine, is designed specifically such that the blade thickness, extending from the trailing edge thickness in the camber line segment toward the center point of the leading edge circle, increases to at least two times and, in particular, to at least 2.5 times the trailing edge thickness.
  • the profile is specially designed such that the above described blade results.
  • FIG. 1 a profile cross section for explanation of the various terms
  • FIG. 2 the course of a blade thickness over a length of a camber line of a profile according to an embodiment of the present invention.
  • FIG. 1 shows, in a partially schematic manner, a profile cross section of a blade element of a guide vane or rotating blade 1 of a gas turbine.
  • the profile has a leading edge (left in FIG. 1 ) and a trailing edge (right in FIG. 1 ), which are connected by a pressure side (top in FIG. 1 ) and an intake side (bottom in FIG. 1 ), as well as a wedge angle ⁇ of 16° at the trailing edge.
  • P refers to a common profile tangent at the leading edge and trailing edge
  • V refers to a leading edge tangent at the leading edge, which is perpendicular to the common profile tangent
  • H refers to a trailing edge tangent at the trailing edge, which is perpendicular to the common profile tangent
  • S refers to a camber line, which extends, at an equal distance from the pressure side and the intake side, from a center point M V of a leading edge circle K V inscribed in the profile, which has the leading edge tangent in common with the profile, to a center point M H of a trailing edge circle analogously inscribed in the profile, which has the trailing edge tangent in common with the profile.
  • the blade thickness proceeding from the trailing edge thickness in the camber line segment toward the center point of the leading edge circle, increases in a strictly monotonic manner.
  • the blade thickness can increase linearly in a region of the camber line segment nearer the trailing edge (right in FIG. 1 ), which starts from the center point of the trailing edge circle and the length of which is at least or at most 50% of the length of the camber line segment, toward the center point of the leading edge circle.
  • the blade thickness in the camber line segment a increases in a monotonically decreasing manner toward the center point of the leading edge circle (toward the left in FIG. 2 ).

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/677,343 2014-04-04 2015-04-02 Gas turbine blade Expired - Fee Related US9869184B2 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP14163476.6 2014-04-04
EP14163476.6A EP2927427A1 (de) 2014-04-04 2014-04-04 Gasturbinenschaufel
EP14163476 2014-04-04

Publications (2)

Publication Number Publication Date
US20150285080A1 US20150285080A1 (en) 2015-10-08
US9869184B2 true US9869184B2 (en) 2018-01-16

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EP (1) EP2927427A1 (de)

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3081751B1 (de) * 2015-04-14 2020-10-21 Ansaldo Energia Switzerland AG Gekühlte turbinenschaufel und verfahren zur herstellung dieser schaufel
EP3231996B1 (de) * 2016-04-11 2020-06-17 Rolls-Royce plc Schaufel für eine axialströmungsmaschine
GB201610783D0 (en) * 2016-06-21 2016-08-03 Rolls Royce Plc Trailing edge ejection cooling
EP3839212A1 (de) * 2019-12-20 2021-06-23 MTU Aero Engines AG Laufschaufel für eine strömungsmaschine
DE102021123281A1 (de) * 2021-09-08 2023-03-09 MTU Aero Engines AG Schaufelblatt für einen Verdichter einer Strömungsmaschine
CN114542207B (zh) * 2022-02-22 2024-08-16 中国航发沈阳发动机研究所 一种涡轮后机匣支板外型面造型设计方法
CN116641915B (zh) * 2022-11-18 2026-02-27 中国航发沈阳发动机研究所 一种宽适应性压气机进口导向器叶型

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2102505A (en) 1981-07-24 1983-02-02 United Technologies Corp Airfoil blade
WO1998059175A1 (de) 1997-06-24 1998-12-30 Siemens Aktiengesellschaft Verdichterschaufel und verwendung einer verdichterschaufel
US20050232778A1 (en) 2004-03-30 2005-10-20 Mitsubishi Fuso Truck And Bus Corporation Blade shape creation program and method
US20080181780A1 (en) * 2006-04-28 2008-07-31 Toyotaka Sonoda Airfoil for axial-flow compressor capable of lowering loss in low Reynolds number region
EP2360377A2 (de) 2010-02-24 2011-08-24 Rolls-Royce plc Kompressortragflügel
WO2012147938A1 (ja) 2011-04-28 2012-11-01 株式会社Ihi タービン翼
US20140373503A1 (en) * 2013-06-21 2014-12-25 Rolls-Royce Plc Method of finishing a blade

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2102505A (en) 1981-07-24 1983-02-02 United Technologies Corp Airfoil blade
WO1998059175A1 (de) 1997-06-24 1998-12-30 Siemens Aktiengesellschaft Verdichterschaufel und verwendung einer verdichterschaufel
US6264429B1 (en) * 1997-06-24 2001-07-24 Siemens Aktiengesellschaft Compressor blade or vane and compressor using a blade or vane
US20050232778A1 (en) 2004-03-30 2005-10-20 Mitsubishi Fuso Truck And Bus Corporation Blade shape creation program and method
US20080181780A1 (en) * 2006-04-28 2008-07-31 Toyotaka Sonoda Airfoil for axial-flow compressor capable of lowering loss in low Reynolds number region
EP2360377A2 (de) 2010-02-24 2011-08-24 Rolls-Royce plc Kompressortragflügel
WO2012147938A1 (ja) 2011-04-28 2012-11-01 株式会社Ihi タービン翼
US20140112795A1 (en) * 2011-04-28 2014-04-24 Ihi Corporation Turbine blade
US20140373503A1 (en) * 2013-06-21 2014-12-25 Rolls-Royce Plc Method of finishing a blade

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US20150285080A1 (en) 2015-10-08
EP2927427A1 (de) 2015-10-07

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